1. Field of the Invention
The present invention relates to the field of aerospace technology, and more particularly to a dual-vortical-flow hybrid rocket engine.
2. Description of Related Art
The traditional dual-vortical-flow hybrid rocket engine uses a single flow channel or multiple flow channels and uses an axial fluid or gas oxidizer injection. With reference to FIG. 1, it shows a schematic drawing of the conventional hybrid rocket engine. A housing 3′ forms with a main body 1′ and an injection port 2′. The main body 1′ disposes with a solid-state fuel 4′. When the rocket engine starts, it injects the oxidizer from the oxidizer injection nozzle 7′ and the oxidizer flow through the solid-state fuel 4′ at a combustion channel, melts with the solid-state fuel, burns with mixing, and ejects from the nozzle 2′ to produce a thrust. By the above design, the vacuum specific impulse is about 250 seconds, which is below its ideal value of more than 300 seconds. The main reason for the low performance is the diffusion flame structure, which causes low mixing and low combustion efficiency. To improve such rocket engine, it requires increasing the combustion channel, which will increase the volume and weight of the rocket and increase the cost.